CN101737198A - Gas-gas nozzle with constriction section - Google Patents
Gas-gas nozzle with constriction section Download PDFInfo
- Publication number
- CN101737198A CN101737198A CN200810226255A CN200810226255A CN101737198A CN 101737198 A CN101737198 A CN 101737198A CN 200810226255 A CN200810226255 A CN 200810226255A CN 200810226255 A CN200810226255 A CN 200810226255A CN 101737198 A CN101737198 A CN 101737198A
- Authority
- CN
- China
- Prior art keywords
- gas
- section
- contraction section
- nozzle
- diameter
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
Abstract
The invention relates to a gas-gas nozzle with a constriction section. A constriction section is additionally arranged in an oxidant channel. The nozzle with the constriction section comprises an inlet section (1), a constriction section (2) and a nozzle outlet section (3). A propellant is used for forming a certain inlet pressure drop at the head part of a combustion chamber to reduce the speed of an oxidant outlet and the pressure drop of a fuel nozzle. The speed of the oxidant outlet is reduced so as to be beneficial to the combustion of the propellant in the combustion chamber, the design of an engine supply system and the improvement of the combustion efficiency of the propellant.
Description
[technical field]
The present invention relates to the liquid propellant rocket engine nozzle, relate in particular to a kind of coaxial shearing gas-gas nozzle of oxygenant band contraction section.
[background technique]
Nozzle is the vitals in the liquid propellant rocket engine, and its function is that propellant agent is introduced the firing chamber, organizes the burning of propellant agent, makes in the mixture fuel and oxidant ratio proper, and component is even along the firing chamber cross-sectional distribution.
Liquid propellant rocket engine is that propellant agent is the motor of liquid, and still, in the liquid propellant rocket engine of some kind, fuel and oxygenant be converted into gaseous state before spraying into firing chamber blending burning, therefore need study gas-gas nozzle technology.Such as in full flow afterburning circulating liquid rocket motor, all the liquid propellant of flow is introduced into oxygen enrichment precombustion chamber and the primary combustion of Fu Ran precombustion chamber, form the oxygen rich fuel gas and the Fu Ran combustion gas of high temperature, two kinds of combustion gas enter main combustion chamber continuation blending afterburning more then, therefore need to use gas-gas nozzle in the main combustion chamber of full flow afterburning cycle engine, promptly fuel of Pen Sheing and oxygenant are the nozzle of gaseous state.
In the time of in liquid propellant sprays into the firing chamber, more little helping more of propellant agent drop accelerates to atomize and evaporation process, improves combustion efficiency.For making the liquid propellant fragmentation become the little drop of diameter, need liquid propellant to spurt into the firing chamber with certain speed, need the oxygen nozzle that higher spray pressure drop is arranged, 15% to 25% of constant pressure is got in the pressure drop of oxidize nozzle usually, and this can increase the burden of motor supply system, improves the sealing and the designing requirement of motor.
Gas propellant has only blending and combustion process in the firing chamber, so oxygenant can get little nozzle velocity, and oxidant injection speed is little, helps increasing the chemical time of oxygenant in the firing chamber, and the combustion efficiency of propellant agent is raise.But, if use the coaxial shearing nozzle of conventional construction, when oxygen jet expansion speed is too small, cause the oxidize nozzle pressure drop too small, the pressure surge meeting of firing chamber is passed to oxygen nozzle head cavity, influences the engine operation performance.
Be well known that, in the oxidize nozzle passage, increase contraction section, make oxygenant higher flow velocity be arranged, because circulation area increases, make the oxidant stream prompt drop low after the contraction section at contraction section.Therefore, in oxidize nozzle, increase contraction section, oxygenant is imported and exported at nozzle formed certain spray pressure drop, prevent that the chamber pressure fluctuation is passed to head cavity; The flow velocity reduction of oxidant outlet place helps improving combustion efficiency.
[summary of the invention]
The present invention has creatively proposed a kind of coaxial shearing gas-gas nozzle with contraction section, and adopting increases contraction section in oxidant channel, make oxygenant import and export into agent one constant pressure drop at nozzle, and the flow velocity of oxidant outlet is lower, helps the burning of propellant agent.
The structure of band contraction section gas-gas nozzle of the present invention is as described below.
Gas-the gas nozzle of band contraction section comprises nozzle entrance short (1), contraction section (2) and jet expansion section (3).The nozzle entrance section is cylindric, and its diameter is 1.5 to 2 times of contraction section diameter; Entrance length 5 is to 10mm; Contraction section and entrance adopt tapering transition, and angle is 90 to 120 degree.By selected oxidize nozzle pressure drop and flow coefficient, determine that by flow formula the diameter of contraction section (2), contraction section Design of length are 5 to 10mm; Contraction section and jet expansion section adopt tapering transition, and angle is 30 to 90 degree.Oxidant outlet section diameter is 1.4 to 2 times of contraction section diameter.The wall thickness of oxidation outlet is 0.5 to 1.5mm.The ratio of oxidant channel length and diameter is greater than 3; Oxygenant internal channel (4) is the right angle with the angle of outer wall (5), helps propellant agent and goes out to form stable recirculating zone at wall, plays the effect of smooth combustion flame.
The present invention makes propellant agent form a constant pressure drop by increase contraction section in oxidant channel, and the chamber pressure fluctuation is transmitted into head cavity; Increase the oxygen exit area of nozzle, reduce oxidant outlet speed, can improve the combustion efficiency of propellant agent.The present invention is based on and be not limited to liquid propellant rocket engine, can also be used for other and relate to device or the equipment that gas-gas blowout is annotated
[description of drawings]
Fig. 1 is band contraction section gas of the present invention-gas nozzle sectional view
[embodiment]
The present invention proposes to be with contraction section gas-gas nozzle first, can prevent that the chamber pressure fluctuation is transmitted into head cavity, and make propellant agent reach high combustion efficiency.Band contraction section gas-gas nozzle and design method thereof are provided, can have guaranteed the burning that propellant agent can be efficient, stable, design method is simply effective.
Band contraction section gas-gas nozzle as shown in Figure 1 comprises nozzle entrance short (1), contraction section (2), outlet section (3) and fuel nozzle (6).The nozzle entrance section is cylindric, and its area is 1.5 to 2 times of contraction section area; Entrance length 5 is to 10mm; Contraction section and entrance adopt tapering transition, and angle is 90 to 120 degree.By selected oxidize nozzle pressure drop and flow coefficient, determine that by flow formula the diameter of contraction section (2), contraction section Design of length are 5 to 10mm; Contraction section and jet expansion section adopt tapering transition, and angle is 30 to 90 degree.Oxidant outlet section diameter is 1.4 to 2 times of contraction section diameter.The wall thickness of outlet section is 0.5 to 1.5mm.The ratio of oxidant channel length and diameter is greater than 3; Oxygenant internal channel (4) is the right angle with the angle of outer wall (5), helps propellant agent and goes out to form stable recirculating zone at wall, plays the effect of smooth combustion flame.Determine the area and the diameter of outer ring fuel nozzle (6) according to area formula.
Claims (5)
1. band contraction section gas-gas nozzle comprises oxidant injection hole and the annular spray-hole of burning, and it is characterized in that: the oxidant injection hole is divided into 3 sections: entrance (1) contraction section (2) and outlet section (3).
2. band contraction section gas-gas nozzle as claimed in claim 1 is characterized in that: contraction section (2) is for cylindric, and its length is 5 to 10mm, and diameter is determined by flow formula.
3. band contraction section gas-gas nozzle as claimed in claim 1 or 2, it is characterized in that: entrance (1) adopts conical surface transition with contraction section (2), the transition angle is 90 to 120 degree, and entrance (1) length is 5 to 10mm, and its area is 1.5 to 2 times of contraction section area.
4. dual concentric gas-gas nozzle as claimed in claim 3 is characterized in that: the ratio of oxidant channel length and diameter is greater than 3, and outlet section and contraction section adopt tapering transition, and angle is 30 to 90 degree.The outlet section diameter is 1.4 to 2 times of contraction section diameter.
5. as the described dual concentric gas-gas nozzle of the arbitrary claim of claim 1-4, it is characterized in that: the wall thickness of outlet section is that the angle of 0.5 to 1.5mm outlet section passage (4) and exterior edge face (5) is 90 degree.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN200810226255A CN101737198A (en) | 2008-11-11 | 2008-11-11 | Gas-gas nozzle with constriction section |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN200810226255A CN101737198A (en) | 2008-11-11 | 2008-11-11 | Gas-gas nozzle with constriction section |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN101737198A true CN101737198A (en) | 2010-06-16 |
Family
ID=42461129
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN200810226255A Pending CN101737198A (en) | 2008-11-11 | 2008-11-11 | Gas-gas nozzle with constriction section |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN101737198A (en) |
Cited By (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103649511A (en) * | 2011-07-07 | 2014-03-19 | 斯奈克玛公司 | Injection element |
| TWI504538B (en) * | 2013-05-31 | 2015-10-21 | Nat Applied Res Laboratories | Dual-vortical-flow hybrid rocket engine |
| CN106461210A (en) * | 2014-06-03 | 2017-02-22 | 西门子公司 | Pumpless metal atomisation and combustion using vacuum generation and suitable material flow control |
| CN110545612A (en) * | 2019-09-04 | 2019-12-06 | 北京航空航天大学 | Multi-stage ionization rotating magnetic field acceleration helicon plasma source |
| CN112196697A (en) * | 2020-10-19 | 2021-01-08 | 北京天兵科技有限公司 | Integrated structure injector for rocket engine |
| CN113432150A (en) * | 2021-06-29 | 2021-09-24 | 中国人民解放军国防科技大学 | Controllable flame stabilizer, engine and aircraft |
| CN113969849A (en) * | 2021-09-26 | 2022-01-25 | 中国人民解放军战略支援部队航天工程大学 | Single-nozzle rocket engine with modular design |
-
2008
- 2008-11-11 CN CN200810226255A patent/CN101737198A/en active Pending
Cited By (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103649511A (en) * | 2011-07-07 | 2014-03-19 | 斯奈克玛公司 | Injection element |
| TWI504538B (en) * | 2013-05-31 | 2015-10-21 | Nat Applied Res Laboratories | Dual-vortical-flow hybrid rocket engine |
| CN106461210A (en) * | 2014-06-03 | 2017-02-22 | 西门子公司 | Pumpless metal atomisation and combustion using vacuum generation and suitable material flow control |
| CN110545612A (en) * | 2019-09-04 | 2019-12-06 | 北京航空航天大学 | Multi-stage ionization rotating magnetic field acceleration helicon plasma source |
| CN112196697A (en) * | 2020-10-19 | 2021-01-08 | 北京天兵科技有限公司 | Integrated structure injector for rocket engine |
| CN113432150A (en) * | 2021-06-29 | 2021-09-24 | 中国人民解放军国防科技大学 | Controllable flame stabilizer, engine and aircraft |
| CN113432150B (en) * | 2021-06-29 | 2022-04-22 | 中国人民解放军国防科技大学 | Controllable flame stabilizer, engine and aircraft |
| CN113969849A (en) * | 2021-09-26 | 2022-01-25 | 中国人民解放军战略支援部队航天工程大学 | Single-nozzle rocket engine with modular design |
| CN113969849B (en) * | 2021-09-26 | 2023-07-11 | 中国人民解放军战略支援部队航天工程大学 | Single-nozzle rocket engine with modularized design |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN101737198A (en) | Gas-gas nozzle with constriction section | |
| US8099940B2 (en) | Low cross-talk gas turbine fuel injector | |
| US11628455B2 (en) | Atomizers | |
| US8545215B2 (en) | Late lean injection injector | |
| US20060059915A1 (en) | Effervescence injector for an aero-mechanical system for injecting air/fuel mixture into a turbomachine combustion chamber | |
| US20120291439A1 (en) | Lean direct fuel injector | |
| JP2007113888A (en) | Combustor structure of gas turbine engine | |
| AU2019325620B2 (en) | Linear throttling high regression rate vortex flow field injection system within a hybrid rocket engine | |
| US8919132B2 (en) | Method of operating a gas turbine engine | |
| RU2011106121A (en) | METHOD AND DEVICE FOR LIQUID FUEL COMBUSTION | |
| US10914237B2 (en) | Airblast injector for a gas turbine engine | |
| US9182124B2 (en) | Gas turbine and fuel injector for the same | |
| US9212609B2 (en) | Combination air assist and pilot gaseous fuel circuit | |
| US11846425B2 (en) | Dual fuel gas turbine engine pilot nozzles | |
| CN105121960A (en) | Injection system for a combustion chamber of a turbine engine, comprising an annular wall having a convergent inner cross-section | |
| CN103175221A (en) | Gas-assisted dual-fuel nozzle used for chemical regenerative cycle | |
| RU2298729C1 (en) | Jet nozzle | |
| CN101737197A (en) | Dual concentric gas-gas nozzle | |
| CN103953465B (en) | A kind of natural gas engine gas rail assembly that is integrated with blender | |
| CN103388530A (en) | Fuel injector with mixing circuit | |
| CN109297018B (en) | Low nitrogen gas combustor | |
| US8381531B2 (en) | Gas turbine fuel injector with a rich catalyst | |
| CN109595551B (en) | Thick-thin burner | |
| JPH0457927B2 (en) | ||
| CN203273901U (en) | Gas nozzle |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| C06 | Publication | ||
| PB01 | Publication | ||
| C10 | Entry into substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| C02 | Deemed withdrawal of patent application after publication (patent law 2001) | ||
| WD01 | Invention patent application deemed withdrawn after publication |
Application publication date: 20100616 |